(1) Field of the Invention
The present invention relates to the field of rotorcraft rotors, and it relates more particularly to devices for damping the movements of blades fitted to such rotors.
The present invention provides a device for damping lead-lag oscillations of the blades of a rotorcraft rotor. More specifically, the present invention relates to ways of mounting a damper included in such a damper device on a hub of the rotor.
(2) Description of Related Art
Rotorcraft are aircraft with a rotary wing, with this classification including helicopters. A helicopter comprises in particular at least one main rotor with a substantially vertical axis that provides the rotorcraft in flight with lift, with propulsion, and with guidance. A helicopter also commonly has a tail rotor for providing it with yaw guidance, and possibly a propulsive propeller in the context of a high-speed and long-range helicopter, commonly referred to as a “hybrid” helicopter.
A rotorcraft rotor comprises a rotary hub driven in rotation by a power plant of the rotorcraft, with blades being mounted on the hub so as to be driven in rotation by the hub. The blades are mounted on the hub so as to be movable in pivoting about their own axes in order to enable the pitch of the blades to be varied at least collectively, as in a tail rotor or in a propulsive propeller, or else otherwise also cyclically as in a main rotor, in particular. Varying the collective or cyclic pitch of the blades of a rotor serves to modify the behavior in flight of the rotorcraft.
It is common practice for the blades to be mounted on the hub via connection members arranged as arms or as sleeves or as other analogous members for mounting a blade on the hub. Such a connection member, referred to below as a “sleeve”, is interposed between the root of a blade and the hub. In order to allow said variations in blade pitch, sleeves are hinged to the hub, e.g. by means of a spherical abutment member, and they are movable in pivoting by a rod for controlling variation in blade pitch.
Since the blades are driven in rotation by the hub and since they are movable in the general plane in which they lie in order to vary their pitch, their individual behavior has the reputation of being complex, particularly for the blades of a main rotor.
It should be considered that the blades are subjected in rotation to forces that vary along their length. When the rotorcraft is hovering or in forward flight, the distribution of aerodynamic forces along a blade gives rise to a distribution of bending moments, and the value of the bending moment is very large at the blade root. When the rotorcraft is moving in translation, the “advancing” blade has an angle of incidence (or pitch) that is smaller than a “retreating” blade for which the pitch is increased in order to balance lifts.
Proposals have therefore been made to hinge the blades on the hub so that they can flap vertically about a flapping axis that is oriented orthogonally relative to the axis of rotation of the hub. When the rotary wing is set into rotation, the combination of centrifugal force and the lift forces causes the blades to tilt with vertical flapping so that the rotary wing takes up a somewhat conical shape, the plane of rotation of the blades being not necessarily the same as a plane orthogonal to the axis of rotation of the hub. In the mounting of a blade on the hub, account must also be taken of a facility for retracting the blades into a folded position of the rotary wing.
In this context, it should also be understood that the blades are also hinged to the hub to perform lead-lag movements in their plane of rotation about a lead-lag axis that is oriented substantially parallel to the axis of rotation of the hub. Such lead-lag hinging of the blades serves to avoid generating bending moments in the blades in their plane.
Nevertheless, the individual oscillations of the blades about their lead-lag axes give rise to a known phenomenon of the rotorcraft presenting ground resonance. Such a phenomenon has the reputation of being potentially dangerous in the event of the resonant frequency of oscillation of the blades about their lead-lag axes coming close to a resonant frequency of the aircraft on the ground. Such a problem arises mainly for the main rotor, however it must also be taken into consideration for other rotary wings of a helicopter, such as for the tail rotor, where consideration needs to be given to the resonant modes of oscillation of the tail boom carrying the tail rotor.
In order to remedy that problem, it is known to fit rotors with damper devices for damping the lead-lag oscillations of the blades about their lead-lag axes. Various damper devices have been developed that make use of elastically deformable dampers. Each damper is in hinged engagement both with a member for engaging the hub via linkage and with a blade that is associated with the damper. Specifically, the linkage comprises one or more links or other analogous elements for transmitting mechanical forces.
The ability of the damper to deform elastically between two fastening points is used in particular for damping the lead-lag oscillations of at least the blade with which it is associated. The damper is placed under elastic deformation stress between said fastening points. One of the fastening points of the damper, considered as being a “distal” fastening point, is anchored via the linkage to a member for engaging the hub. The other of said fastening points of the damper, considered as being a “proximal” fastening point, is engaged with said blade that is associated therewith.
In various possible configurations, the member for engaging the damper with the hub is arranged on a neighboring blade or is incorporated in the hub. The engagement of the damper at its proximal fastening point with a said blade with which it is associated may potentially take place via the sleeve carrying the blade. Still among the various possible configurations of greater or lesser complexity seeking to ensure that the damper is stressed in a manner that matches requirements, the distal fastening point of the damper, and possibly also its proximal fastening points, are in hinged engagement with a link. Such a link may itself potentially be hinged to other links that are hinged to one another and/or to the hub and/or possibly to the sleeve of a blade.
The dampers used may be arranged in various ways. For example, the dampers may be of elongate shape and work in compression/traction, or they may be of cylindrical shape and work in twisting. A damper of cylindrical shape presents the advantage of having an arrangement that is compact and also presents the advantage of being easy to install in a sleeve used for connecting the blade to the hub.
By way of example, proposals are made in document FR 2 943 621 (Eurocopter SAS), to house a lead-lag damper of cylindrical shape inside a said sleeve. The proximal fastening point of the damper is fastened to the sleeve, while its distal fastening point is connected via a link to a said engagement member incorporated in the hub.
By way of example, reference may also be made to the following documents: FR 2 653 405 (Aerospaciale Société Nationale Industrielle SA); FR 2 733 961 (Eurocopter France SA); and EP 2 223 854 (Agusta SPA), those documents relating to installing dampers on a rotor in various configurations of greater or lesser complexity.
It has been found that installing dampers inside the sleeves, as disclosed in document FR 2 943 621 (Eurocopter SAS), gives rise to various other advantages. Dampers that are installed outside sleeves give rise to aerodynamic drag that it is desirable to avoid. Furthermore, installing dampers inside the sleeves makes it possible to protect them from the hostile environment of the rotor.
Nevertheless, it has also been found that installing dampers inside the sleeves is made easier when the dampers are of cylindrical shape. Such a damper can easily be anchored inside a sleeve and can be put into hinged engagement with the hub via a said link. Such an installation is difficult to arrange for a damper of any structure.
It is thus found that the solution proposed in document FR 2 943 621 (Eurocopter SAS) is applicable for a rotorcraft of given structure and as a function of specific needs for damping lead-lag oscillations of blades. It is difficult or even inappropriate to transpose the use of that solution with any rotorcraft that might present a variety of structures.
As mentioned above, the requirements for damping lead-lag oscillations of blades are closely associated with said phenomena of resonance, and therefore with the power and the structure specific to the rotorcraft. It is appropriate for the installation of the damper in the sleeve to avoid constituting any obstacle to optimizing potential configurations for installing other members of the damper device on the rotor. It would be advantageous for various linkage configurations adapted to the specific requirements for damping lead-lag oscillations of the blades to be made potentially available without requiring any major structural modification to the damper device.